Known gas turbine engines incorporate shaft-mounted turbine blades circumferentially circumscribed by a turbine casing or housing. Hot gasses flowing past the turbine blades cause blade rotation that converts thermal energy within the hot gasses to mechanical work, which is available for powering rotating machinery, such as an electrical generator. Referring to FIGS. 1-4, known turbine engines, such as the gas turbine engine 30 include a multi stage compressor section 32, a combustor section 34, a multi stage turbine section 36 and an exhaust system 38. Atmospheric pressure intake air is drawn into the compressor section 32 generally in the direction of the flow arrows F along the axial length of the turbine engine 30. The intake air is progressively pressurized in the compressor section 32 by rows rotating compressor blades and directed by mating compressor vanes to the combustor section 34, where it is mixed with fuel and ignited. The ignited fuel/air mixture, now under greater pressure and temperature than the original intake air, is directed to the sequential rows R1, R2, etc., in the turbine section 36. The engine's rotor and shaft 39 has a plurality of rows of airfoil cross sectional shaped turbine blades 40 terminating in distal blade squealer tips 46 in the compressor 32 and turbine 36 sections. For convenience and brevity further discussion of turbine blades and abradable layers in the engine will focus on the turbine section 36 embodiments and applications, though similar constructions are applicable for the compressor section 32. Each blade 40 has a concave profile pressure side 42 and a convex suction side 44. The high temperature and pressure combustion gas, flowing in the combustion flow direction F imparts rotational motion on the blades 40, spinning the rotor 39s. As is well known, some of the mechanical power imparted on the rotor shaft is available for performing useful work. The combustion gasses are constrained radially distal the rotor by turbine casing 60 and proximal the rotor by air seals. Referring to the Row 1 section shown in FIG. 2, and the perspective view of the same blade 40 in FIG. 3, respective upstream vanes 62 direct upstream combustion gases generally parallel to the incident angle of the leading edge 48 of turbine blade and downstream vanes redirect downstream combustion gas exiting the trailing edge 50 of the blade.
The turbine engine 30 turbine casing 60 proximal the blade squealer tips 46 is lined with a plurality of sector shaped abradable components 64, each having a support surface retained within and coupled to the casing 60 and an abradable substrate 66 that is in opposed, spaced relationship with the blade tip by a blade tip gap G. The abradable substrate is often constructed of a metallic/ceramic material that has high thermal and thermal erosion resistance and that maintains structural integrity at high combustion temperatures. As the abradable surface 66 metallic-ceramic materials is often more abrasive than the turbine blade tip 46 material a blade tip gap G is maintained to avoid contact between the two opposed components that might at best cause premature blade tip wear and in worse case circumstances might cause engine damage.
In addition to the desire to prevent blade tip 46 premature wear or contact with the abradable substrate 66, for ideal airflow and power efficiency each respective blade tip 46 desirably has a uniform blade tip gap G relative to the abradable component 64 that is as small as possible (ideally zero clearance) to minimize blade tip airflow leakage L between the concave pressure blade side 42 and the convex suction blade side 44 as well as axially in the combustion flow direction F. However, manufacturing and operational tradeoffs require blade tip gaps G greater than zero. Such tradeoffs include tolerance stacking of interacting components, so that a blade constructed on the higher end of acceptable radial length tolerance and an abradable component abradable substrate 66 constructed on the lower end of acceptable radial tolerance do not impact each other excessively during operation. Similarly, small mechanical alignment variances during engine assembly can cause local variations in the blade tip gap G. For example in a turbine engine of many meters axial length, having a turbine casing abradable substrate 66 inner diameter of multiple meters, very small mechanical alignment variances can impart local blade tip gap G variances of a few millimeters.
During turbine engine 30 operation the turbine engine casing 60 may experience out of round (e.g., egg shaped) thermal distortion. Casing 60 thermal distortion potential increases between operational cycles of the turbine engine 30 as the engine is fired up to generate power and subsequently cooled for servicing after thousands of hours of power generation. Commonly, greater casing 60 and abradable component 64 distortion tends to occur at the uppermost and lowermost casing circumferential positions (i.e., 6:00 and 12:00 positions) compared to the lateral right and left circumferential positions (i.e., 3:00 and 9:00). For example, if casing distortion at the 6:00 position causes blade tip contact with the abradable substrate 66 one or more of the blade tip squealers 46 may be worn during operation, increasing the blade tip gap locally in various other less deformed circumferential portions of the turbine casing 60 from the ideal gap G to a larger gap. The excessive blade gap distortion increases blade tip leakage L, diverting hot combustion gas away from the turbine blade 40 airfoil, reducing the turbine engine's efficiency.
The exemplary blade 40 squealer tip 46 construction and its interaction with the turbine casing abradable surface 66 is shown in greater detail in FIGS. 3-6. The squealer tip 46 has a an airfoil planform tip plate 56 having along its outer periphery downstream from its leading edge 48 and upstream from its trailing edge 50 opposed and laterally separated outwardly or radially projecting concave pressure 52 and convex suction 54 rails, which respectively have opposed inner faces and outer faces. An enclosed tip cavity 57 is defined between the tip plate 56 and respective inner faces of the pressure rail 52 (also referenced in FIG. 4 as the pressure rail inner surface 53) and suction rail 54 from the leading 48 to trailing 50 edges. Referring to the streamline simulation of gas flow between and around the squealer tip 46 and the abradable surface (the abradable surface is not shown for clearer flow streamline viewing), pressure side gas flow FP is deflected around the leading edge 48 and separates from contact with the pressure side rail 52, allowing heat to concentrate on the outer face of the pressure rail. Such excessive heat concentration can cause pressure rail 52 erosion, prematurely wearing out the blade and undesirably increasing the blade tip gap, as previously described. Combustion gas flow FT undesirably passes through the blade tip gap over the top of the squealer tip 46, but most of it is diverted away from the pressure rail inner surface 53 toward the suction side rail, creating another potential heat concentration zone along the pressure rail inner surface. Gas flow FS along the suction side 44 of the blade tip 46 is directed toward the blade trailing edge 50, where it cannot assist in transfer of heat from the pressure rail 52 heat concentration zone. As previously mentioned, friction contact between the squealer tip 46 pressure rail 52 and the abradable surface 46 also undesirably increases pressure rail area heat concentration.
Another known conventional blade squealer tip 146 is shown in FIG. 7, having a segmented pressure side rail 152 with a slot 158 proximal the squealer tip 146 trailing edge 150. In this embodiment the suction side rail 154 is continuous downstream from the leading edge 148 to the trailing edge 150. The rails 152, 154 and the underlying tip plate (not shown) form the squealer tip cavity 157.